U.S. patent number 6,102,335 [Application Number 09/074,772] was granted by the patent office on 2000-08-15 for elliptical orbit satellite, system, and deployment with controllable coverage characteristics.
This patent grant is currently assigned to Mobile Communications Holdings, Inc.. Invention is credited to Jay Brosius, David Castiel, John Draim, Matthew Schor.
United States Patent |
6,102,335 |
Castiel , et al. |
August 15, 2000 |
Elliptical orbit satellite, system, and deployment with
controllable coverage characteristics
Abstract
An elliptical orbit satellite system which describes
communication and TT&C with ground stations. Earth stations are
located for the circular orbiting satellite in a way such that the
line of sight can never include geo synchronous satellites. The
ground stations for the elliptically orbiting satellites monitor
the position of the satellite, and no antenna communicates with a
satellite which is in direct line of sight between the antenna and
a geo satellite. Another aspect of the invention locates two
TT&C stations, separated in longitude by 90.degree. and
configures these stations such that each satellite in each
constellation will be able to communicate with one of the two
TT&C stations once during each satellite rotation period.
Inventors: |
Castiel; David (Washington,
DC), Draim; John (Washington, DC), Brosius; Jay
(Washington, DC), Schor; Matthew (Washington, DC) |
Assignee: |
Mobile Communications Holdings,
Inc. (Washington, DC)
|
Family
ID: |
26892707 |
Appl.
No.: |
09/074,772 |
Filed: |
May 8, 1998 |
Related U.S. Patent Documents
|
|
|
|
|
|
|
Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
|
197260 |
Feb 16, 1994 |
5582367 |
|
|
|
892239 |
Jun 2, 1992 |
5931417 |
|
|
|
Current U.S.
Class: |
244/158.4;
455/430 |
Current CPC
Class: |
B64G
1/1007 (20130101); H04B 7/195 (20130101); B64G
1/242 (20130101); H04B 7/18576 (20130101); B64G
1/1085 (20130101); B64G 1/002 (20130101) |
Current International
Class: |
B64G
1/00 (20060101); B64G 1/24 (20060101); B64G
1/10 (20060101); H04B 7/195 (20060101); H04B
7/185 (20060101); B64G 001/00 () |
Field of
Search: |
;244/158R ;455/12.1,13.1
;342/356 ;453/427-430 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Barefoot; Galen L.
Attorney, Agent or Firm: Fish & Richardson P.C.
Parent Case Text
This is a continuation-in-part of application Ser. No. 08/197,260
filed Feb. 16, 1994, U.S. Pat. No. 5,582,367, which is a
continuation-in-part of application Ser. No. 07/892,239 filed Jun.
2, 1992, U.S. Pat. No. 5,931,417.
Claims
What is claimed is:
1. A method of interfacing with orbiting satellites, which requires
communicating with each of a plurality of satellites in a plurality
of orbiting rings at least once during each rotation,
comprising:
locating a plurality of satellites which orbit the earth in two
elliptical rings having opposite inclinations to one another;
establishing telemetry stations for said plurality of satellites at
two locations on the earth with a 90.degree. separation in
longitude therebetween; and
using said two stations, and no others, to view each of the
satellites in said rings at least once during each satellite
rotation.
2. A method as in claim 1 wherein said earth stations are located
between around 25.degree. and 40.degree. north latitude.
3. A method as in claim 1 comprising the further step of orbiting a
plurality of low earth orbit circular or slightly elliptical
satellites in an equatorial orbit, and communicating with said
circular satellites using said same two stations.
4. A satellite communications system comprising:
a constellation of satellites orbiting the earth including two
rings of oppositely-inclined elliptical satellites each having a
plurality of satellites therein, and a ring of circular equatorial
low earth orbit satellites;
two and only two tracking telemetry and control stations, located
on the earth, and communicating with said satellites, said two
stations separated in longitude by 90.degree., and communicating
with said satellites such that each satellite in all of said orbits
communicates with at least one of said stations once during each of
its rotation periods.
5. An apparatus as in claim 4 wherein said elliptical satellites
have an inclination with an absolute value of around 116.degree.,
an eccentricity of about 0.35, and a semi-major axis of about
10,000 kilometers.
6. An apparatus as in claim 5 wherein said circular orbits have a
height above the earth of 7800 kilometers.
7. An apparatus as in claim 5, wherein at least some of said
satellites have apogees which always point in a constant direction
relative to the sun.
Description
FIELD OF THE INVENTION
The present invention relates to elliptical satellite orbits,
constellations, methods, and communication systems. The present
invention also explores specific earth station placement optimized
for these elliptical orbit constellations, however, it could also
apply tc other types of constellations including weather,
surveillance, and the like.
BACKGROUND OF THE INVENTION
The concept of artificial satellites circling the earth was
introduced to scientific literature by Sir Isaac Newton in 1686.
Things have gotten considerably more complicated since that time,
however. The basic concepts of an orbit are described in any
orbital mechanics or astrodynamics textbook, such as "Fundamentals
of Astrodynamics" by Bate et al. or "Orbital Mechanics" by
Chobotov, AIAA Education Series, Publisher. The following
definitions of these terms will be first provided here, since they
are necessary for proper understanding of the present
invention.
The earliest satellites placed into space by man were deployed into
very low circular orbits. The resulting visibility footprint of one
of these satellites was quite small and a single satellite had the
added disadvantage of providing only a few minutes of coverage per
day. In fact, it was quite common for an observer on the equator to
miss being in contact with such a satellite for several days.
Raising the satellite to a higher orbital altitude (e.g.,
.apprxeq.600 nautical miles) helped extend both the coverage
footprint, average viewing elevation, and the time in view, but for
some missions frequent or even continuous coverage became a
requirement. This led to the deployment of early multiple satellite
systems, a typical example being the Navy's Transit navigation
satellite system. Satellite systems designers were increasingly
asked to provide continuous coverage; first, for latitudinal zones
and then, for the entire globe.
One of the first constellation designers to study zonal coverage
was David Luders. The Englishman, John Walker, was the first to
systematize the design of multiple-ring, multiple satellites per
ring, constellations and his work contributed greatly to the
optimization of a number of multi-satellite systems (e.g., NAVSTAR
GPS). A Russian designer, G. Mozhaev, independently came up with
similar arrays using a more theoretical approach based on
mathematical set and group theory. Polar constellations often
employed the concept of "street-of-coverage", and further coverage
improvements were made by Beste, Ballard and Rider. More recently,
Hanson and Linden have investigated large arrays of low earth orbit
"LEO" satellites (40-200 satellites). All of these designers
employed circular orbits; and even with this simplification,
constellation design was considered at best a difficult and time
consuming trial and error exercise.
The motion of any artificial satellite may be described using a
number of parameters. The eccentricity, e, is a measure of the
amount of ellipticity. An orbit which has a greater eccentricity
number is more elliptical. Eccentricity e=0 would describe a
circle, any number between 0 and 1 is an ellipse, and the
eccentricity number of 1 or greater would be a parabola or a
hyperbola, respectively (curves which never close).
For an elliptical orbit, the earth, or the object being orbited, is
at one of the focal points of the ellipse. Therefore, the satellite
is sometimes closer to the earth than at other times. The apogee is
defined as the point of highest altitude of a satellite, while
perigee is the point of lowest altitude.
A retrograde orbit is one in which the direction of revolution is
opposite to that of the earth. A posigrade or prograde orbit is an
orbit in which the satellite revolves around the earth in the same
direction as the earth.
The inclination angle i is an angle measured between the plane of
the orbit, and a plane of the reference, usually the Equator. An
inclination angle i less than 900 is a prograde orbit, while an
inclination angle greater than 90.degree. is a retrograde orbit. A
90.degree. orbit is a polar orbit.
The period, T, is a measure of how long the satellite takes to make
one entire orbit. Mean anomaly M is another way to describe the
position in the orbit. Mean anomaly is a fictitious angle
indicating the fraction of 360 degrees corresponding to the
fraction of the period through which the satellite has passed at
any point of its orbit.
The Right Ascension of the Ascending Node ("RAAN") is an angle
between the first point of Aries (.gamma.), a non-rotating
celestial reference, and the line of nodes, which is the line
forming the intersection of a plane of the orbit and the plane of
the equator. The line of nodes gives a measure of the position or
orientation of the orbit. The longitude of the ascending node n is
the angle between the i unit vector (pointing towards the Greenwich
meridian) and the ascending node, in the rotating reference.
The argument of perigee .omega. is an angle measured in the plane
of the orbit between the point of the ascending node and the
nearest point of perigee.
Most practical satellites prior to the invention by the present
inventors used relatively simple systems based on circular orbits.
The earth was covered symmetrically by multiple satellites, which
each operate to cover a section of the earth.
Elliptical orbits have been typically avoided in the art, because
of their asymmetries, and the consequent problems that they might
cause. However, some individual elliptical orbits and elliptical
orbit constellations have been proposed. The Russian Molniya orbit
is a posigrade orbit designed for polar and high latitude coverage.
Other posigrade orbits have been described by John Draim in his
U.S. Pat. Nos. 4,809,935 and 4,854,527. 4,809,935 describes a
three-satellite constellation giving continuous coverage of the
entire Northern hemisphere, and an extension of this constellation
to include an equatorial orbit resulting in a four-satellite array
giving continuous global coverage of both hemispheres. This latter
four satellite array provided somewhat higher elevation coverage in
the Northern hemisphere than in the Southern Hemisphere.
U.S. Pat. No. 4,854,527 describes a common-period four-satellite
array giving continuous global coverage with satellites at a lower
altitude range than in the first patent. A discussion of obtaining
extra Northern Hemisphere coverage through use of elliptic
satellite constellations may be found in ANSER Space Systems
Division Note SpSDN 84-1, "Satellite Constellation Design
Techniques for Future Space Systems" dated September 1984, by John
Draim and James Cooper. Another application of posigrade elliptic
orbits is the ACE 20 and ACE-Prime orbits developed by Mr. A.
Turner of Loral Corporation.
The present invention also simplifies the design of the solar
panels by requiring no more than 1 or 2 degrees of freedom. In the
example orbit discussed herein which is 116.degree. retrograde, the
panels need only one degree of freedom since the satellite in the
plane containing the
earth-sun line, i.e. midnight or noon ascending node. In a similar
way, a satellite usually needs to radiate its heat toward cold,
empty space. In the present invention, it is much easier to face
the satellite in a way that always faces the heat radiators away
from the sun.
It is also well known that the earth is not totally spherical, but
actually it is rather oblate. That is, the earth is bigger at the
bottom than it is at the top. The J.sub.2 harmonic, due to the
earth's oblateness, causes the node .OMEGA. and argument of perigee
.omega. of an orbit to change. The gravitational pull of the
earth's equatorial bulge causes, for example, the orbital plane of
an eastbound satellite to swing westward. More generally, the force
component is directed towards the Equator. This resultant
acceleration causes any satellite to reach the Equator (node) short
of the crossing point where it would have reached it on a spherical
earth. For each revolution, therefore, the orbit regresses a
.DELTA. amount. These effects have been the subjects of various
attempts at compensation.
Sun synchronous circular orbits are also known. These are orbits
where the rotation rate of the right ascension of the ascending
node is equal to and in the same direction as, the right ascension
rate of the mean sun.
SUMMARY OF THE INVENTION
The previous specifications, of which this is a
continuation-in-part, described the invention of non-uniform
capacity distribution tailored by latitude and population. This was
done using an elliptical satellite array, and tailoring the
parameters by latitude and population, and/or by time of day.
Specific ways of carrying out these options by using a retrograde
inclined orbit, and/or a sun synchronous apogee elliptical orbit
are described.
The present specification also considers the special ground-control
problems posed by these satellite constellations, and more
specifically how to most effectively carry out ground station
communication with ground communication stations ("GCS") and
tracking, telemetry and control ("TT&C") stations.
The satellites of the present invention most preferably communicate
with ground-based TT&C stations at least once during every
revolution of the satellite. This preferred satellite architecture
uses satellites that circle the earth once every three hours.
According to one preferred satellite constellation and orbit, the
satellites orbit at a different rate than that at which the earth
rotates. Therefore, the satellites see different positions on the
earth during different times in their rotation periods. This
produces special constraints on earth station placement. The
inventors of the present invention have discovered a way to
minimize the number of TT&C earth stations which are necessary
to meet the above goal.
According to another aspect of the invention, a low circular orbit
is used in conjunction with the satellites of the present
invention, and to avoid interference with other systems.
Most geosynchronous-orbiting satellites communicate with the earth
in the so-called C-band. Each satellite has a specific
geosynchronous position, to avoid interference between
communications between the various satellites. The inventors
noticed that one problem with low circular orbiting satellites is
that their communications could interfere, at times, with the
geosynchronous satellites. The present invention teaches a way to
avoid this problem by placement of earth stations and by tracking
of positions of various satellites to avoid the interference.
BRIEF DESCRIPTION OF THE DRAWINGS
These and other aspects of the invention will now be described in
detail with reference to the accompanying drawings, wherein:
FIG. 1 shows a first design space for elliptical sun synchronous
retrograde orbits according to a first embodiment of the present
invention;
FIG. 2 shows the characteristics of a special orbit according to a
second embodiment of the present invention in which the apogee is
always pointing towards the sun;
FIG. 3 shows a design space for this second embodiment of the
present invention using prograde orbits;
FIG. 4 shows a constellation of satellites, each orbiting and
communicating with earth stations on the earth;
FIG. 5 shows a rocket and inertial guidance unit used according to
the present invention to propel the rocket into orbit;
FIGS. 6, 7A, 7B, 8A, 8B, 9A, 9B, 9C, 10A and 10B show
characteristics of preferred orbits of the present invention.
FIG. 11 shows a diagram of the satellites orbiting the earth
according to a preferred configuration of the present
invention;
FIG. 12 shows a conceptual diagram of the communication between a
geosynchronous satellite and a low earth orbit circular
satellite;
FIG. 13 shows an azimuth/elevation plot showing the view of a
satellite track from a position at 10.degree. north latitude;
FIG. 14A shows a diagram of a constellation of elliptical inclined
satellite orbits;
FIG. 14B shows communication between an earth station and one of
those elliptical satellites;
FIG. 15 shows communication between that same earth station and a
different one of the elliptical satellites;
FIG. 16 shows a block diagram of control systems in the earth
station computer; and
FIG. 17 shows a conceptual diagram of positioning of earth stations
relative to the satellites.
DESCRIPTION OF THE PREFERRED EMBODIMENT
The present invention exploits the gravitational effects from the
earth's oblateness, in combination with a preferably elliptical
orbit, to allow preferential coverage of different parts of the
earth as a function of parameters which are related to satellite
demand. This has significant advantages since it allows
preferential coverage based on a chosen characteristic, here either
one hemisphere over the other, or time of day.
For instance, a satellite system primarily intended for use cover
the United States would prefer to preferentially cover the Northern
hemisphere as opposed to the Southern hemisphere. More
specifically, by choosing elliptical orbits such that anything
above 40.degree. south latitude was covered, a great majority of
the world's land mass could be covered without wasted capacity.
This embodiment of the invention optimizes the characteristics of
the elliptical satellite to have desired coverage characteristics.
According to this first preferred mode, structure is described for
putting a satellite in a special orbit which preferentially covers
part of the earth over the other part.
The first type of orbits, discussed according to the present
invention herein, are elliptical retrograde orbits which provide
preferential coverage of one part of the earth over the other part
through adjustment of orbital parameters.
As mentioned above, all orbits are effected by the earth's J.sub.2
gravitational term. This term effects the .omega. and .OMEGA. terms
of every orbit. In order to compensate the orbit, the general
equation ##EQU1## must be satisfied. This first embodiment takes a
special case of the equation (1).
The significance of the constant on the right hand of the equality
sign in Equation (1) lies in its synchronism with the Earth's
yearly motion about the Sun. In order to preserve the orientation
of the orbital plane with respect to the earth-sun line, it is
necessary to advance the plane of the orbit by 360 degrees/365.25
days or 0.9856 deg/day.
Specifically, the effect of J.sub.2 term on .OMEGA. and .omega. can
be expressed as follows: ##EQU2## , where n is the mean motion in
degrees per day, R.sub.e is the earth's equatorial radius, a is the
semi major axis in kilometers, e is the eccentricity, i is the
inclination and the change in .OMEGA. and .omega. are both in
degrees per day.
According to this first embodiment, we want to make d.omega./dt
term approach zero. Luckily, this can be easily done by adjustment
of the sine term in equation 3 to zero. Therefore, we set 5
sin.sup.2 i=4, requiring that sin.sup.2 i=4/5 or i=arc sin {square
root (4/5)}; so i=63.435.degree. or its complement
116.565.degree..
This embodiment preferably uses an elliptical orbit of 116.565
degrees. The prior art has used circular sun synchronous orbits.
All so-called circular orbits may have some slight degree of
ellipticity. For purposes of this specification, an elliptical
orbit is defined as an orbit whose ellipticity is greater than
0.002. This effectively excludes circular orbits which are slightly
elliptical due to imperfections in the orbits. These elliptical
orbits, with e.apprxeq.=0.001 are sometimes called frozen
orbits.
Therefore, we set ##EQU3## to zero, and we set ##EQU4## to +0.9856,
the amount per day by which the earth+s revolves around the sun. By
substituting this into equation (3), a set of combinations of
apogee, perigee and inclination are found which satisfy the
attached formula which are shown in the attached FIG. 1.
For an elliptical sun synchronous orbit, only a very small
circumscribed part of this design space can be used. First, this
satellite should have no apsidal rotation, to keep the apogee in
one hemisphere. Accordingly, the inclination must be
116.565.degree.. A certain amount of leeway is possible, however,
and practically speaking the orbit can be inclined anywhere between
115 and 118.degree. and still obtain sufficiently stable
characteristics, although some minor orbit corrections may be
necessary from time to time.
Along this line, only a certain class of orbits are usable.
Circular orbits are known in the prior art, and do not have the
ability to produce the preferential coverage characteristics in the
way done according to the present invention. Therefore, a leftmost
limit on the design space shown by point 102 in FIG. 1 represents
the limit to require an elliptical orbit. The rightmost limit is
set by the minimum satellite height at perigee. A satellite orbit
should be, practically speaking, greater than, for example, 100
nautical miles. Preferably, the lowest limit is 250. The point 104
represents the position where perigee will fall below 100 nautical
miles. Therefore, the design space extends between the points 102
and 104. Within this design space, the inclination varies between
115.degree. and 119 .degree.. The usable design area is therefore
shown in the box in FIG. 1.
Within that box, period varies from 2.6 to 3.1 hours, apogee varies
from 100 to 4600 nautical miles, and perigee varies from 100 to
2200 nautical miles.
These orbits allow the coverage to be adjusted, or biased, to favor
the Northern hemisphere over the Southern hemisphere.
More specifically, the allowable range of orbital parameters
includes orbital periods between 2.68 and 3.1 hours, and orbital
eccentricities between 0.002 and 0.38.
The postulated orbit preferably has an orbit or orbits with the
integral period value of 3.0 hours. This 3-hour orbit with
corresponding mean motion of an even 8 revolutions per day will
result in a repeating ground track. The use of other, non-integral
values for orbital period(s) still results in the satellite's
ground track crossing the Equator on the ascending and descending
nodes at given values of local time, but the points of such
crossings will not now occur at fixed longitudinal points. Any
point along the design space horizontal line (116.565 degrees) may
be selected to provide a base line set of orbital parameters upon
which such an orbit or constellation may be configured.
Applications
This invention may be used for communications, earth sensing,
surveillance, weather, or any other satellite function found useful
for satisfying mission requirements. The invention can be used in a
single satellite mode, and will provide better coverage during
daylight hours than during nighttime hours. Effectively, coverage
is "stolen" from nighttime coverage and diverted to daytime
coverage. The most probable future application of the invention in
this case will be found in the construction and use of ordered
arrays (or constellations) of such satellites.
In order to show how this system would be used, a few examples from
the design space in FIG. 1 will be discussed herein. These examples
are analyzed using a computer program such as Orbital Workbench, or
OSAC written by the Naval Research Lab, or Graftrak, available from
Silicon Solutions, Inc; Houston Tex. This program is run with the
inclination, apogee and other information from the chart in FIG. 1.
The characteristics of that orbit are obtained. Then, the desired
characteristics are used to modify the orbit until the proper
places from the design space are identified. Some preferred orbits
according to the present invention will be described herein.
The second embodiment of the present invention is one which
produces a special kind of elliptical orbit. This special orbit has
a constant-pointing apogee, which faces in a constant direction
relative to the sun all year round. This is obtained by a posigrade
orbit in which the equation ##EQU5## is satisfied.
FIG. 2 shows a resulting sun synchronous orbit with apogee pointing
towards the sun. This preferred embodiment of the present invention
comprises a satellite in an orbit which has a sun synchronous
apogee which assumes an orbit around the earth such that the apogee
of the satellite is always facing towards the sun. The satellite
100 is shown with its orbit 102, orbiting the earth 104. Different
seasons find the earth at different portions around the sun, and
these portions are shown as positions 110, 120, 130 and 140. The
apogee point, shown as element 142, is always facing the sun.
To obtain the preferred operating range for this equation, the
equations
These orbits have characteristics which are synchronous with
respect to the time of year. By specifying any initial RAAN and
epoch, therefore, the Right Ascension of the apogee of this orbit
will stay constant over time with respect to the sun. For one
special class of orbits, the apogee will always be pointing towards
the sun as shown in FIG. 3. For another special class of orbits,
the apogee will be pointing for example at 2 degrees relative to
the sun. In any of these orbits, therefore, the apogee is
controlled to be constant.
For this embodiment, the apogee is always at a constant right
ascension angle from the right ascension of the earth-sun line:
usable inclinations range from 0 to 43 degrees, usable periods from
1.7 to 5.0 hours (again, preferably 3 hours to obtain a repeating
ground track), and usable eccentricities from 0.0002 to 0.56.
A few examples of how these orbits would be chosen and the
characteristics thereof are explained herein.
According to another preferred mode of the invention, the first
and/or second embodiments are further modified to include multiple
satellite configurations. This modification comprises a
constellation of satellites which preferentially cover the Northern
hemisphere, as compared with the
Southern hemisphere or vice versa.
The constellation of satellites orbiting the earth 400 is shown in
FIG. 4. Of course, it should be understood that while FIG. 4 shows
only three satellites, 402, 403 and 404, in reality there would be
many more. These two satellites are located and operate to
preferentially cover one portion of the earth over another (first
embodiment) and/or one time of day (second embodiment) over
another.
Each of the satellites communicates with a earth-based earth
station, shown schematically as station 406, in a conventional way
to exchange information therewith. Accordingly, the present
invention also contemplates use of an earth station with such
satellites, this earth station having characteristics to track
satellites having the characteristics discussed above, and to
communicate therewith. There are a plurality of earth stations,
each. positioned on the earth, and each including tracking
equipment to track a motion of at least one of said satellites.
Each earth station, and each satellite also includes communication
equipment to communicate between the earth station and the at least
one satellite.
The satellites according to the present invention are initially
boosted into their orbits by special rockets of the type intended
to deliver satellites. One such rocket, 500, with the satellite 502
therein is shown in FIG. 5. The rocket includes a first stage
engine 504, of any known solid or liquid fuel type, and a second
stage engine 506. Rocket engines are well known in the art, and it
will be assumed that the second stage engine is a liquid type
rocket fuel engine. This engine combines a liquid fuel with an
oxygenator at point 508, which ignites the fuel. The ignition
accelerates the speed of the fuel through a constriction 510,
causing a sonic shock wave shown as 512 which travels out the
output nozzle 514. (It must be understood that the fixture in FIG.
5 shows this stage rocket with the first stage still attached.)
The rocket is controllable both in direction and in thrust. More
generally, the vector control of the rocket is controllable.
The rocket is controlled by an onboard navigation computer 516. The
basic characteristics of a booster rocket and guidance unit are
shown, for example, in U.S. Pat. No. 4,964,340, the disclosure of
which is herewith incorporated by reference.
According to a fourth embodiment of the rocket of the present
invention, the inertial guidance unit is controlled to boost the
rocket into an elliptical retrograde orbit selected from the design
space box around line 100 shown in FIG. 1. The satellite is then
delivered into that orbit, to maintain that orbit.
According to a fifth embodiment of the present invention, the
rocket of FIG. 5 has an internal guidance unit which is programmed
to boost the rocket into a posigrade orbit of an elliptical type,
selected from the design space shown in FIG. 3. At that time, the
satellite is released into the orbit, to thereby maintain
thereafter the appropriate orbit.
The third, fourth and fifth embodiments are usable in combination
with either of the first or second embodiments described above.
Some examples of the preferred orbits used according to the present
invention will now be described.
First preferred orbit configuration
The first preferred orbit is a four satellite minimum array ring
which covers any northern hemisphere region north of 20.degree.
north latitude during daylight hours, with a minimum 15 degree
elevation angle .sigma.. The satellites have an optimized afternoon
ascending node, a three hour period and an argument of perigee
.omega. other than 270. The ellipse actually therefore tilts
towards the sun and provides a ring of orbits which are both sun
synchronous and always have their apogee pointing towards the
sun.
The characteristics of these orbits are such that the satellites
appear to be moving backwards from west to east since they are in
retrograde orbit.
Using the basic satellites discussed above, selection of the main
orbital parameters were adjusted through trial iterations beginning
around the beginning values of .omega.=270 and RAAN=F(YY, MM, YY,
HH, MM, and SS). The resulting graph track view show visibility
circles and lines which reach down to a certain latitude.
This system is very unique, since with only four LEO-MEO
satellites, all regions north of 20.degree. latitude can be covered
with visibility angles of 15.degree.. It would take three to four
times as many circular satellites to do the same thing.
Second preferred orbit configuration
The second preferred orbit covers everything in the northern
hemisphere above 20.degree. north latitude both day and night. One
ring of satellites has noon ascending nodes and the other has
midnight ascending nodes. This has the significant advantage of
simplifying the design of the solar array of the satellite.
Most satellites have solar arrays, which need to face the sun in
order to power the satellite. If we use an orbit like the present
example, then this solar array needs only one degree of freedom to
follow the sun. This simplifies the satellite design. This
requirement is satisfied by placing one ring with noon ascending
nodes and another ring with RAANs displaced 180.degree. from the
first ring and having midnight ascending nodes.
FIGS. 6, 7A and 7B show this basic orbit. FIG. 6 shows the noon
orbit, and the four satellites therein, labelled 01, 02, 03, and
04. FIG. 7A shows the midnight ring, with the satellites labelled
05, 06, 07, and 08. FIG. 7B shows the noon plus midnight rings. The
combined view of FIG. 7B shows that most of the coverage is in the
northern hemisphere. There is only spotty coverage in the southern
hemisphere, but the clustering is in the north.
Third preferred orbit configuration
A third example is a six satellite equatorial, prograde, apogee
pointing towards the sun orbit. This third example uses terms of
the formula for advance of the line of nodes at 0.9856.degree. per
day and provides an extra degree of redundancy and higher elevation
angles in the tropical and equatorial zones.
Fourth preferred orbit configuration
The fourth example is another equatorial prograde orbit with apogee
pointing towards the sun with only four satellites. This array
emphasizes continuous equatorial region daytime coverage with
visibility angle of 10.degree.. FIG. 8A shows 1100 GMT which is
daylight over Europe, and shows that most of Europe is well
covered. However, Europe is less well covered at 2300 GMT shown in
FIG. 8B.
Fifth preferred orbit configuration
The fifth preferred orbit constrains the visibility angle to
0.degree. and obtains continuous equatorial region daytime coverage
with only three satellites. Again, there are gaps at nighttime, but
none in the daylight hours. FIGS. 9A, 9B, and 9C show the various
daylight hour coverages. FIG. 9A shows coverage at 1535 GMT, FIG.
9B shows coverage at 740 GMT, and FIG. 9C shows coverage at 1500
GMT.
Sixth preferred orbit configuration
The sixth preferred orbit is shown in FIGS. 10A and 10B. This four
satellite array combines classic sun synchronism condition of
##EQU7## with the apogee on the sunward side of the earth.
.omega.=262.degree.. Apogee always occurs close to the meridian of
the Earth at local apparent noon. This four satellite array
provides continuous coverage of day lit areas north of 20.degree.
north latitude all year round in all countries. This sixth example
has an afternoon ascending node, apogee at noon, a forced inclined
plane, and a three hour period with the apogee equals about 4000
nautical miles.
Preferred Constellation and Earth Station Placement
The preferred satellite constellation which covers most of the
earth, but preferentially covers the northern hemisphere, is shown
in FIG. 11. This constellation includes two inclined elliptical
rings of satellites 1100 and 1102. Each of these inclined
elliptical satellite rings preferably includes four satellites,
each having perigees in the southern hemisphere. The present
inventors have called these Borealis(TM) orbits. These orbits have
a period of three hours, a semi major axis of 10,561 kilometers,
eccentricity of 0.347, inclination of 116.5.degree., right
ascension of the ascending node of 89.198 and argument of perigee
of 269.6. Since the orbit period is 3 hours, the satellites repeat
their ground track every 24 hours. Rings 100 and 102 each include
four satellites, equally spaced in time and hence having different
mean anomalies. It should be understood that the descriptions given
herein are for the ring 1102, and that the ring 1100 has very
similar but complementary characteristics.
The low earth orbit ring 1104 is a circular or moderately-elliptic
earth orbit with a 7800 kilometer height above the earth, 0.degree.
inclination and which preferably includes eight satellites.
This constellation communicates with ground control stations. FIG.
12 shows a ground control station 1200 at a location on the equator
intended to track a geosynchronous ("geo") satellite. Looking from
this hypothetical point on the equator, there would be times when
the extension of the beam to the geo satellite and the beam to the
low earth orbit circular satellite would cross. At that moment, the
feeder uplink would be illuminating both satellites and both
satellites might re-transmit the signals to the ground. FIG. 12
shows if communication with satellite 1106 were attempted from
position 1200, this could cause interference with communications to
the geo satellite 1202 if both communications used the same
band.
The present technique avoids that interference, and therefore
allows use of standard and commercially-available satellite
communication equipment. All of this equipment operates on the same
band. The present inventors, by determining a way in which
interference could be avoided, make it commercially feasible to use
satellite orbits which would otherwise potentially interfere with
geosynchronous satellites.
The inventors of the present invention noticed that if an earth
station were located at point 1204, the ground view of the
satellites would be that shown in FIG. 13. Specifically, by
locating the GCS station at least at 5.degree. north latitude, and
preferably at 10.degree. north latitude to 40.degree. north
latitude, this problem may be avoided.
A typical feeder beam 3.degree. in diameter 1300 is shown tracking
the satellite 1106. This beam never intersects the geo satellite
1202. Looking from the ground, in fact, an observer would see a
line 1302 for the equatorial satellites in low earth orbit, and
would see the line 1304 for the satellites in geo orbit. So long as
these paths remain separated by at least the earth station beam
width, interference is avoided. This allows the low earth satellite
orbit to use the same band of communication which is used for a geo
satellite.
According to this aspect of the present invention, therefore, the
ground station is situated at a location where the low earth orbit
satellite will always appear in the sky at a position which is
lower than locations of equatorial geo satellites, by an amount of
separation at least equal to a beam width. Preferably this
separation is 0.5.degree., and the beam width is 3.degree..
The above interference problem is caused by communication with the
equatorial satellite LEO ring. A second interference problem is
caused when communicating with the inclined elliptical satellite
rings 1100 and 1102. Each elliptical satellite in the inclined
orbit, at some point in the course of its orbit, will cross the
path of a geo satellite as seen by an observer at a specific point
on the earth. This is because the earth station is at a point on
the earth. It is inevitable that sometimes a straight line will be
formed which includes that point on the earth, an elliptical
satellite, and a geo satellite. The elliptical satellite
communications beam would then have the geo satellite within its
path.
The geo satellite operators would object to any system which could
cause a possibility of interference with their communications.
However, since C band communications equipment is commercially
available, it would be most feasible to use this equipment.
The inventors of the present invention noticed that the Borealis
orbits 1100 and 1102 communicate with the largest communications
area at the apogee of their orbits. At points in the satellite's
rotation where the altitude is below heights of around 4000 km down
to perigee, less preferably 4000 to 5000 km down to perigee, the
amount of land mass with which the satellite communicates is
minimal.
FIGS. 14A and 14B show the geometry of the satellite ring 1102.
Ground control station 1400 is shown communicating with satellite
b4. Satellite b4 produces a footprint 1402 on the earth. The
satellite constellations are moving counterclockwise in the sense
of FIG. 14B. As the satellite b4 approaches its perigee, however,
the feeder beam would impinge on the area of the ring of
geostationary satellites.
According to the present invention, the ground station 1400
includes at least two antennas, 1404 and 1406 shown in FIG. 16. As
the satellite b4 approaches its minimum operational altitude, the
30 wide feeder link beam 1405 illuminates an area which includes
the geosynchronous satellite ring, as shown in FIG. 14B. When the
link beam approaches that area, the feeder beam from the antenna
1404 is switched off, and the feeder beam from antenna 1406, which
tracks the satellite b3, is switched on as shown in FIG. 15. This
feeder link beam is now servicing satellite b3, and hence has no
possibility of interference with the geosynchronous satellites.
The ground station 1400 should be located at a latitude of
45.degree. north or less to make this operation most effective.
FIG. 16 shows a more detailed layout of the earth station 1400. The
antenna 1404 directs its beam 1510 towards the satellite b4. At the
same time, the position of the antenna 1404 is being controlled by
appropriate commands sent to motor 1512 which controls the azimuth
and elevation of the antenna. The motors are also producing
signals, coupled through signal lines to a navigation and control
computer 1514. Navigation computer monitors these signals, and from
the signals calculate the position in space of the beam 1405
illumination, and extension of that position to the height of geo
satellites. The navigation computer processes this position
information to determine if the 3.degree. beam, once extended to
the height of geo satellites, will intersect the area of geo
satellites. If so, navigation computer terminates the transmission
from antenna 1404, and commands antenna 1406 to begin transmitting
beam 1520 to satellite b3. Antenna 1404 is then moved to track
satellite b2 (not shown), to prepare for the moment when satellite
b3 will reach a position that would intersect the field of geo
satellites. This technique avoids any possibility of communications
interference with the elliptical satellites.
Yet another aspect of the invention involves the placement of
telemetry, tracking and control stations ("TT&C stations").
FIG. 11 shows an earth station, which we can assume to be a
TT&C station for purposes of this embodiment, at point A at
noon (or midnight). The earth station at position A can communicate
with both rings 1100 and 1102. Therefore, each satellite will be
able to communicate with the earth station A, at least at one point
during each rotation period, at 12:00 noon and at midnight. This
communication will be possible for three hours before and after
noon or midnight. Therefore, from approximately 9:00 AM--3:00 PM
and 9:00 PM--3:00 AM all satellites can communicate with earth
station A, which will be able to carry out TT&C for all
satellites in the constellation.
At 6:00 AM and 6:00 PM, the earth has rotated such that the same
earth station is at position B. That position will be able to
communicate with the rings 1102, but the satellites in the ring
1100 will be over the horizon relative to position B and therefore
it will not be possible to communicate from position B to satellite
ring 1100.
The present inventors have devised a way to communicate with every
satellite in every ring using only two TT&C stations. FIG. 17
shows TT&C stations 1700 and 1702. These stations are at
position A at 12:00; and station 1700 communicates with all
satellites. At 3:00 PM, the earth station locations are shown as B.
At this time, station 1700 communicates with satellite ring 1102,
and station 1702 communicates with satellite ring 1100.
FIG. 17B shows the positions at 6:00 (position C), and positions at
9:00 (position D). The positions at 12:00 are again position A,
shown in FIG. 17.
The situation continues throughout the entire earth rotation, as
the TT&C stations pass in and out of various satellite
orbits.
The inventors of the present invention have found that by locating
TT&C stations at 25.degree.-35.degree. north latitude, and
separated by 90.degree. of longitude, that two TT&C stations
can service the entire Concordia/Borealis constellation. This means
that two TT&C ground stations can be used to service all
satellites in the system, including those in inclined planes and
those in the equatorial plane.
Preferably, the present inventors postulate locating one of the
TT&C stations in the U.S., and the other off of northern
Spain.
Although only a few embodiments have been described in detail
above, those having ordinary skill in the art will certainly
understand that many modifications are possible in the preferred
embodiment without departing from the teachings thereof.
All such modifications are intended to be encompassed within the
following claims.
* * * * *